Staged gas turbine combustion chamber with counter swirling arrays of radial vanes having interjacent fuel injection

ABSTRACT

Gas turbine engine combustion chamber has staged combustion to reduce nitrous oxides and includes a first radial flow swirler and a second radial flow swirler located axially of an annular mixing zone with each swirler having vanes for rotating the incoming air in substantially opposite directions relative to each other; first and second fuel injectors are provided with a first fuel injectors located in one of the passages of each of the first and second swirlers and with the second fuel injectors located upstream of the passages of the first and second swirlers.

The present invention relates to a gas turbine combustion chamber, andto a method of operating a gas turbine engine combustion chamber.

In order to meet emission level requirements for industrial low emissiongas turbine engines, the engine combustion chamber volumes has beenincreased. Currently industrial gas turbine engines use annular orcan-annular combustion chambers. The requirement to increase the volumeof the combustion chamber assembly whilst incorporating the combustionchamber assembly in the same axial length has necessitated the use of aplurality of tubular combustion chambers, whose longitudinal axes arearranged in generally radial directions. The inlets of the tubularcombustion chambers are at their radially outer ends, and transitionducts connect the outlets of the tubular combustion chambers with a rowof nozzle guide vanes to discharge the hot exhaust gases axially intothe turbine sections of the gas turbine engine.

Also in order to meet the emission level requirements, staged combustionis required in order to minimise the quantity of the oxides of nitrogen(NOx) produced. Currently the emission level requirement is for lessthan 25 volumetric parts per million of NOx for an industrial gasturbine exhaust. The fundamental way to reduce emissions of nitrogenoxides is to reduce the combustion reaction temperature, and thisrequires premixing of the fuel and all the combustion air beforecombustion takes place. The oxides of nitrogen (NOx) are commonlyreduced by a method which uses two stages of fuel injection. Our UKpatent no. 489339 discloses two stages of fuel injection to reduce NOx.In staged combustion, both stages of combustion seek to provide leancombustion and hence the low combustion temperatures required tominimise NOx. The term lean combustion means combustion of fuel in airwhere the fuel to air ratio is low i.e. less than the stoichiometricratio.

The present invention seeks to provide a novel gas turbine combustionchamber, and a novel method of operating a gas turbine engine combustionchamber.

Accordingly the present invention provides a gas turbine enginecombustion chamber comprising first air intake means, primary fuelinjector means and a first fuel and air mixing zone, the first fuel andair mixing zone being defined by at least one annular wall and anupstream wall connected to the upstream end of the annular wall, theupstream wall having at least one aperture, the first air intake meanscomprising at least one first radial flow swirler and at least onesecond radial flow swirler, each first radial flow swirler beingarranged to supply air into the first fuel and air mixing zone throughsaid aperture, each second radial flow swirler being arranged to supplyair into the first fuel and air mixing zone through said aperture, eachfirst radial flow swirler being positioned axially downstream of therespective second radial flow swirler with respect to the axis of thecombustion chamber, each first radial flow swirler being arranged toswirl air in the opposite direction to the respective second radial flowswirler, the primary fuel injector means being arranged to supply fuelinto at least one of the passages defined between the vanes of each ofthe first radial flow swirlers and into at least one of the passagesdefined between the vanes of each of the second radial flow swirlers.

Preferably at least one pilot fuel injector is provided, each pilot fuelinjector is aligned with a respective one of the apertures to supplyfuel into the first fuel and air mixing zone.

Preferably the primary fuel injector means is arranged to supply fuelinto all the passages defined between the vanes of the first radial flowswirler.

Preferably the primary fuel injector means is arranged to supply fuelinto all the passages defined between the vanes of the second radialflow swirler.

Preferably the primary fuel injector means is arranged to supply fuelinto the radially outer region of the passages between the vanes.

The primary fuel injector means may comprise a hollow cylindrical memberarranged to extend axially with respect to the combustion chamber, thecylindrical member has a plurality of apertures spaced apart axiallyalong the cylindrical member to inject fuel into the passages.

The apertures may be arranged to direct the fuel radially inwardly.

The primary fuel injector means may be arranged to inject gas fuel orevaporated liquid fuel.

The pilot fuel injector may be arranged to inject gas fuel, or liquidfuel.

The combustion chamber may be tubular and has a single aperture in itsupstream wall.

The combustion chamber may further comprise secondary air intake means,secondary fuel injector means and a secondary fuel and air mixing zone,the secondary fuel and air mixing zone is annular and surrounds thefirst fuel and air mixing zone, the secondary fuel and air mixing zonebeing defined at its radially outer extremity by a second annular wall,the secondary fuel injector means being arranged to supply fuel into theupstream end of the secondary fuel and air mixing zone, the secondaryfuel and air mixing zone being an fluid flow communication at itsdownstream end with the interior of the combustion chamber downstream ofthe first fuel and air mixing zone.

The secondary air intake may be downstream of the first air intakemeans.

The secondary fuel and air mixing zone may be defined at its radiallyinner extremity by a third annular wall.

The annular wall may have a first portion defining the first fuel andair mixing zone, a second portion of increased diameter downstream ofthe first portion and a third frusto conical portion interconnecting thefirst and second portions.

The third conical portion may have a plurality of equi-circumferentiallyspaced apertures arranged to direct the secondary fuel and mixture fromthe secondary fuel and air mixing zone as a plurality of jets in adownstream direction towards the centre line of the combustion chamber.

The apertures may be slots.

The downstream end of the second annular wall may be secured to thethird conical portion of the annular wall.

Cooling air may be supplied to an annular chamber defined between theannular wall and the third annular wall.

The secondary fuel injector means may comprise a plurality ofequi-circumferentially spaced fuel injectors.

The secondary fuel injector means may be arranged to inject gas fuel orevaporated liquid fuel.

The downstream end of the first portion of the annular wall reduces indiameter to a throat.

The combustion chamber may comprise tertiary air intake means, tertiaryfuel injector means and a tertiary fuel and air mixing zone, thetertiary fuel and air mixing zone is annular and surrounds the secondarycombustion zone, the tertiary fuel and air mixing zone is defined at itsradially outer extremity by a fourth annular wall, the tertiary fuelinjector means is arranged to supply fuel into the upstream end of thetertiary fuel and air mixing zone, the tertiary fuel and air mixing zoneis in fluid flow communication at its downstream end with a tertiarycombustion zone in the interior of the combustion chamber downstream ofthe secondary combustion zone.

The annular wall may have a fourth portion of larger diameter than thesecond portion downstream of the second portion and defining thetertiary combustion zone, a fifth frusto conical portion interconnectingthe second and fourth portions.

The downstream end of the second portion of the annular wall may reducein diameter to a throat.

The tertiary air intake may be downstream of the second air intakemeans.

The tertiary fuel and air mixing zone may be defined at its radiallyinner extremity by a fifth annular wall.

The fifth conical portion may have a plurality of equi-circumferentiallyspaced apertures arranged to direct the tertiary fuel and air mixturefrom the tertiary fuel and air mixing zone as a plurality of Jets in adownstream direction towards the centreline of the combustion chamber.

The apertures may be slots.

The downstream end of the fourth annular wall may be secured to thefifth conical portion of the annular wall.

The tertiary fuel injector means may comprise a plurality ofequi-circumferentially spaced fuel injectors.

The tertiary fuel injectors means may be arranged to inject gas fuel orevaporated liquid fuel.

Fuel may only be supplied from the pilot fuel injector into the firstfuel and air mixing zone from the start of operation of the gas turbineengine until a predetermined output power level is obtained, fuel issupplied from the primary fuel injector means into at least one of thepassages defined between the vanes of the first radial flow swirler andinto at least one of the passages defined between the vanes of thesecond radial flow swirler to flow into the first fuel and air mixingzone for output power levels greater than the predetermined level, andsimultaneously fuel is supplied from the secondary fuel injector meansinto the secondary fuel and air mixing zone to flow into the interior ofthe combustion chamber downstream of the first fuel and air mixing zone.

Fuel may be supplied from the pilot fuel injector only into the firstfuel and air mixing zone from the start of operation of the gas turbineengine until a predetermined output power level is obtained, supplyingfuel from the primary fuel injector means into at least one of thepassages defined between the vanes of the first radial flow swirler andinto at least one of the passages defined between the vanes of thesecond radial flow swirler to flow into the first fuel and air mixingzone for output power levels greater than a predetermined level, andsimultaneously supplying fuel into the secondary fuel and air mixingzone to flow into the secondary combustion zone in the interior of thecombustion chamber downstream of the first fuel and air mixing zone,supplying fuel into the tertiary fuel and air mixing zone to flow intothe tertiary combustion zone in the interior of the combustion chamberdownstream of the secondary combustion zone for output power levelsgreater than a second predetermined level and for ambient airtemperatures greater than a predetermined temperature.

The predetermined output power level may be 35 to 40% power.

The present invention will be more fully described by way of examplewith reference to the accompanying drawings, in which:

FIG. 1 is a view of a gas turbine engine having a combustion chamberassembly and fuel injector according to the present invention.

FIG. 2 is an enlarged longitudinal cross-sectional view through thecombustion chamber shown in FIG. 1.

FIG. 3 is a further enlarged longitudinal cross-sectional view throughthe upstream end of the combustion chamber assembly shown in FIG. 2.

FIG. 4 is a cross-section in the direction of arrows G--G in FIG. 3, and

FIG. 5 is a cross-sectional view in the direction of arrows H--H in FIG.3.

FIG. 6 is a graph of percentage base load fuel flow versus percentageload for the combustion chamber shown in FIG. 3.

FIG. 7 is an enlarged longitudinal cross-sectional view through theupstream end of an alternative combustion chamber assembly according tothe present invention.

FIG. 8 is an enlarged longitudinal cross-sectional view through theupstream end of a further combustion chamber assembly according to thepresent invention.

An industrial gas turbine engine 10, shown in FIG. 1, comprises in axialflow series an inlet 12, a compressor section 14, a combustion chamberassembly 16, a turbine section 18, a power turbine section 20 and anexhaust 22. The turbine section 18 is arranged to drive the compressorsection 14 via one or more shafts (not shown). The power turbine section20 is arranged to drive an electrical generator 26, via a shaft 24.However the power turbine section 20 may be arranged to provide drivefor other purposes. The operation of the gas turbine engine 10 is quiteconventional, and will not be discussed further.

The combustion chamber assembly 16 is shown more clearly in FIGS. 2 to5. A plurality of compressor outlet guide vanes 28 are provided at theaxially downstream end of the compressor section 14, to which is securedat their radially inner ends an inner annular wall 30 which defines theinner surface of an annular chamber 34. A diffuser is defined between anannular wall 32 and the upstream portion of the inner annular wall 30.The downstream end of the inner annular wall 30 is secured to theradially inner ends of a row of nozzle guide vanes 90 which direct hotgases from the combustion chamber assembly 16 into the turbine section18.

The combustion chamber assembly 16 comprises a plurality of equallycircumferentially spaced tubular combustion chambers 36. The axes of thetubular combustion chambers 36 are arranged to extend in generallyradial directions. The inlets of the tubular combustion chambers 36 areat their radially outermost ends and their outlets are at their radiallyinnermost ends.

Each of the tubular combustion chambers 36 comprises an upstream wall 44secured to the upstream end of an annular wall 37. A first, upstream,portion 38 of the annular wall 37 defines a first fuel and air mixingzone 64, and a second, downstream portion 42 of the annular wall isinterconnected with the first portion 38 by a third portion 40. Thesecond portion 42 of the annular wall has a greater diameter than thefirst portion 38, and the third portion 40 is frusto conical.

A plurality of equally circumferentially spaced transition ducts 46 areprovided, and each of the transition ducts 46 has a circularcross-section at its upstream end. The upstream end of each of thetransition ducts 46 is located coaxially around the downstream end of acorresponding one of the tubular combustion chambers 36, and each of thetransition ducts 46 connects and seals with an angular section of thenozzle guide vanes 90.

A plurality of cylindrical casings 48 are provided, and each cylindricalcasing 48 is located coaxially around a respective one of the tubularcombustion chambers 36. Each cylindrical casing 48 is secured to arespective boss 52 on an annular engine casing 50. A number of chambers54 are formed between each tubular combustion chamber 36 and itsrespective cylindrical casing 48.

The upstream end of each transition duct 46 has a bracket 56 whichextends radially, with respect to the upstream end of the transitionduct, and the engine casing 50 has a plurality of pairs of brackets 58.Each bracket 56 is pivotally secured to a respective one of the pairs ofbrackets 58 by a pin 60, to provide a pivot mounting which is describedmore fully in our copending UK patent application no. 9019089.3 filedSept. 1, 1990.

The upstream wall 44 of each of the tubular combustion chambers 36 hasan aperture 62 to allow the supply of air and fuel into the first fueland air mixing zone 64. A plurality of first radial flow swirlers areprovided and each first radial flow swirler is arranged coaxially withthe aperture 62 in the upstream wall 44 of the respective tubularcombustion chamber 36. Similarly a plurality of second radial flowswirlers are provided and each second radial flow swirler is arrangedcoaxially with the aperture 62 in the upstream wall 44 of the respectivetubular combustion chamber 36. The first radial flow swirlers arepositioned axially downstream, with respect to the axis of the tubularcombustion chamber, of the second radial flow swirlers.

Each first radial flow swirler comprises a first side plate 66, a secondside plate 68 and a plurality of first vanes 70. The first side plate 66has a central aperture arranged coaxially with the aperture 62 in theupstream wall 44, and the plate 66 is secured to the upstream wall 44.The first vanes 70 extend axially between and are secured to the firstand second side plates 66 and 68 respectively. A number of passages 72are formed between the first vanes 70 for the flow of air. Each secondradial flow swirler comprises a plurality of second vanes 74 and a thirdside plate 76. The second vanes 74 extend axially between the secondside plate 68 and the third side plate 76. The second side plate 68 hasa central aperture arranged coaxially with the aperture 62 in theupstream wall 44, and has a shaped annular lip 78 which extends in anaxially downstream direction into the aperture 62. A number of passages80 are formed between the second vanes 74 for the flow of air. The firstand second vanes 70,74 of the first and second radial flow swirlers arearranged to swirl air in opposite directions, as seen from FIGS. 4 and5. A first annular air intake 82 is defined axially between the radiallyouter end of each first side plate 66 and a closure plate 84 at theouter end of each cylindrical casing 48.

A plurality of pilot fuel injectors 86 are provided, and each pilot fuelinjector 86 is arranged coaxially with the aperture 62 of one of thetubular combustion chambers 36 to supply fuel through the aperture 62into the first fuel and air mixing zone 64. A plurality of primary fuelinjectors 88 are provided for each of the tubular combustion chambers36. Each of the primary fuel injectors 88 comprises a hollow cylindricalmember which extends axially with respect to the tubular combustionchamber 36. Each of the hollow cylindrical members passes axiallythrough the third side plate 76 and the second side plate 68 and locatesin a blind hole in the first side plate 66. Each of the hollowcylindrical members is arranged to pass axially through one of thepassages 80 between the second vanes 74 and through one of the passages72 between the first vanes 70. The hollow cylindrical members arepositioned towards the radially outer region of the passages 72,80, andhave axially spaced apertures 90 to inject fuel into the first radialflow swirler assembly and axially spaced apertures 92 to inject fuelinto the second radial flow swirler assembly. The apertures 90 and 92are arranged to direct the fuel radially inwardly.

A second annular fuel and air mixing zone 94 surrounds the first fueland air mixing zone 64 of each tubular combustion chamber 36. Eachsecond annular fuel and air mixing zone 94 is defined between a secondannular wall 96 and a third annular wall 98. The second annular wall 96defines the radially outer extremity of the second fuel and air mixingzone 94 and the third annular wall 98 defines the radially innerextremity of the second fuel and air mixing zone 94. The axiallyupstream end 100 of each third annular wall 98 is secured to the firstside plate 66 of the first radial flow swirler of the respective tubularcombustion chamber 36. A second annular air intake 102 is definedaxially between the upstream end of each second annular wall 96 and theupstream end 100 of the respective third annular wall 98 to supply airinto the second annular fuel and mixing zones 94.

A plurality of secondary fuel injectors 104 are provided for each of thetubular combustion chambers 36. Each of the secondary fuel injectors 104comprises a hollow cylindrical member which extends axially with respectto the tubular combustion chamber 36. Each of the hollow cylindricalmembers passes axially through the upstream end 100 of the third annularwall 98 to supply fuel into the second fuel and air mixing zone 94.

Each second and third annular wall 96,98 is arranged coaxially aroundthe first portion 38 of the respective annular wall. At the downstreamend of each second annular fuel and air mixing zone 94, the second andthird annular walls 96 and 98 are secured to the respective third frustoconical portion 40, and each frusto conical portion 40 is provided witha plurality of circumferentially spaced apertures 106 which are arrangedto direct fuel and air into a second combustion zone 112 in the tubularcombustion chambers 36, in a downstream direction towards the axis ofthe tubular combustion chamber 36. The apertures 106 may be circular orslots.

Each first side plate 66 is provided with a plurality of apertures 108to supply cooling air into an annular space 110 between the upstreamportion 38 of the annular wall and the third annular wall 98 for coolingof the annular wall.

The annular wall may be formed from a laminated structure comprisingspaced perforated inner and outer sheets which give transpirationcooling of the annular wall.

In operation primary air A flows through the first air intake 82 andthrough the first and second radial flow swirlers. The lips 78 directthe primary air into the first fuel and air mixing zone or primarycombustion zone 64. The flows of air from the first and second radialflow swirlers are in opposite directions and this produces opposed flowvortices B and C. A shear layer D is formed between the vortices B and Cwhich improves mixing turbulence.

The pilot injectors 86 only are used at low power settings, that is lessthan about 40% power. They inject gas or pre-evaporated liquid fuel at anarrow angle only into the primary air which has passed through thesecond radial flow swirlers to create a locally fuel rich mixture on theaxes of the tubular combustion chambers 36. Diffusion causes the fuel tomix with the primary air in the vortex B. Vortex C remains an air onlyregion. Thus a locally fuel rich mixture is created on the combustionchamber 36 centreline which sustains combustion in the primarycombustion zone 64.

The primary fuel injectors 88 are not used, during low power operation,and thus only primary air exits from the downstream end of the passages72 and 80 formed between the respective vanes 70 and 74 of the first andsecond swirler assemblies.

At high power settings, at or greater than about 40% power, the pilotinjectors 86 are not used, and all the fuel supplied into the combustionchamber 36 is supplied from the primary and secondary injectors 88 and104 respectively.

At high power settings, at or greater than 40% power, the primary fuelinjectors 88 inject gas, or pre-evaporated liquid fuel, into thepassages 72 and 80 formed between the respective vanes 70 and 74 of thefirst and second swirler assemblies. Simultaneously the secondary fuelinjectors 104 inject gas, or pre-evaporated liquid fuel, into the secondfuel and air mixing zone 94 to mix with secondary air entering thesecond fuel and air mixing zone 94 through the second annular intake102.

The first and second radial flow swirler assemblies direct the fuel andair mixture towards the centreline of the tubular combustion chamber 86before it is turned so that it flows parallel to the centreline of thecombustion chamber 36. The fuel is entrained into both vortex B andvortex C which have opposite swirl, and the shear layer D between thetwo vortices improves the mixing turbulence. There is no net swirl inthe tubular combustion chamber 36 and therefore the gases diffusequickly back to the centreline of the tubular combustion chamber 36 inprimary combustion zone 64 enabling the volume of the tubular combustionchamber 36 to be minimised and also minimising mixing with cooling airon the inner surface of the upstream portion 38 of the combustionchamber 36. This minimises heat transfer to the upstream portion 38 ofthe combustion chamber, allows more efficient use of the cooling air andthus improves combustion efficiency.

Secondary air E flows through the second air intake 102 into thesecondary air and fuel mixing zone 94. The secondary air and fuel ismixed as it flows axially downstream through the second fuel and airmixing zone 94. The resulting fuel and air mixture formed in thesecondary air and fuel mixing zone 94 is injected through the apertures106 into the second downstream portion 42 of the tubular combustionchamber 36 where secondary combustion occurs in the second combustionzone 112. The fuel and air mixture injected from the second fuel and airmixing zone 94 is in the form of discrete Jets F which are directed in adownstream direction towards the centreline of the tubular combustionchamber 36. This ensures good penetration of the secondary fuel and airmixture into the gases from the primary combustion zone 64 and hencegood mixing. Interaction of the secondary fuel and air mixture Jets Fwith cooling air flowing over the inner surfaces of the downstreamportion 42 of the combustion chamber 36 is minimised because of thisangling of the jets F towards the centreline of the combustion chamber.

The graph in FIG. 6 illustrates how the fuel flow to the pilot, primaryand secondary injectors 86,88 and 104 respectively varies with thepower, or load, setting of the gas turbine engine.

Only the pilot injectors 86 are supplied with fuel at power settingsbelow 35% power. At power settings above 35% fuel is suppliedsimultaneously to the primary and secondary injectors 88 and 104, andthe supply of fuel to the pilot injectors 86 is terminated. At a power,or load setting of 35%, 83% of the fuel supplied to each combustionchandler is supplied to the primary injectors 88 and the remaining ofthe fuel is supplied to the secondary injectors 104. As the power, orload, setting is increased the total quantity of fuel supplied to eachcombustor increases and the total quantity of fuel supplied to theprimary injectors and secondary injectors increases. The percentage ofthe total fuel supplied to the combustion chamber, which is supplied tothe primary injectors 88 decreases gradually from 83% at 35% powersetting to approximately 50% at 100% power setting. The percentage ofthe total fuel supplied to the combustion chamber which is supplied tothe secondary injectors 104 increases gradually from 17% at 35% powersetting to approximately 50% at 100% power setting.

The percentage of fuel supplied to the primary injectors 88 preferablydecreases gradually from 28% at 40% power setting to 50% at 100% powersetting whilst the percentage of fuel supplied to the secondaryinjectors 104 increases from 22% at 40% power setting to 50% at 100%power setting.

The first fuel and air mixing zone 64 is supplied with fuel so that ithas a constant maximum temperature of 1800° K. (1527° C.) to preventdisassociation of nitrogen at higher temperatures, and hence prevent theformation of NOx.

The second combustion zone 112 is supplied with fuel so that it also hasa constant maximum temperature of 1800° K. (1527° C.),and has a minimumtemperature of 1500° K. (1227° C.) to prevent the build up of carbonmonoxide etc. Preferably the mimimum temperature is 1550° K. The heatliberated in the first fuel and air mixing zone 64 heats the secondaryair in the second fuel and air mixing zone 94.

In the combustion chamber 36 shown in FIGS. 2 to 5, it is required thatthe temperature of the flame in the first fuel and air mixing zone 64remains substantially constant, or within a predetermined range oftemperatures, so that the emissions of NOx remains low. However, withvariations of power setting between 35% and 100% power, the marginbetween the required flame temperature and the temperature at which theflame is extinguished varies. In some circumstances the flame may beextinguished in the first fuel and air mixing zone. In order to providean adequate margin between the flame temperature and the temperature atwhich the flame is extinguished, a greater proportion of fuel could besupplied to the first fuel and air mixing zone 64. However, thissolution is not desirable because the flame temperature is increased andthus the emissions of NOx is increased.

An alternative combustion chamber assembly 136, shown in FIG. 7, issubstantially the same as that shown in FIGS. 2 to 5 and the samereference numerals have been used to designate like parts. Thecombustion chamber assembly 136 differs from that shown in FIGS. 2 to 5in that the downstream end of the first portion 38 of the annular wall37 has a frusto conical portion 120 which reduces in diameter to athroat 122. The third frustoconical portion 40 interconnects the firstportion 38 and the second portion 42, and the second portion 42 stillhas a greater diameter than the first portion 38.

The reduction in diameter at the downstream end of the first portion 38,provided by the frustoconical portion 120 and the throat 112, enhancesthe recirculation of hot combustion products into the first fuel and airmixing zone, or primary combustion zone 64, to reignite the fuel and airmixture. This, it is believed, also minimises or prevents secondary airflowing from the second fuel and air mixing zone 94 into the first fueland air mixing zone 64 or primary combustion zone. The reduction indiameter at the downstream end of the first portion 38, in combinationwith a constant temperature in the first fuel and air mixing zone orcombustion zone 64 allows a suitable margin between the flametemperature and the temperature at which the flame is extinguished to bemaintained with variations of power setting between 35% and 100% powerto prevent the flame in the first fuel and air mixing zone 64 beingextinguished.

The fuel flows to the pilot, primary and secondary injectors 86,88 and104 respectively varies with the power setting of the gas turbine enginein the same manner as that illustrated in FIG. 6.

The combustion chambers shown in FIGS. 2 to 5 and in FIG. 7 are suitablefor operation across the full power range for ambient air temperaturesin the range of -30° C. to +30° C. or higher.

A further combustion chamber assembly 236, shown in FIG. 8, is similarto that shown in FIG. 7 and the same reference numerals have been usedto designate like parts. The combustion chamber assembly 236 differsfrom that shown in FIG. 7 in that each of the tubular combustionchambers 236 also comprises a fourth portion 130 positioned downstreamof and interconnected to, the second portion 42 by a fifth portion 132.The fourth portion 130 of the annular wall has a greater diameter thanthe second portion 40, and the fifth portion 132 is frustoconical. Thedownstream end of the second portion 42 of the annular wall 37 has afrustoconical portion 134 which reduces in diameter to a throat 136.

A third annular fuel and air mixing zone 138 surrounds the secondcombustion zone 112 of each tubular combustion chamber 236. Each thirdannular fuel and air mixing zone 138 is defined between a fourth annularwall 140 and a fifth annular wall 142. The fourth annular wall 140defines the radially outer extremity of the third fuel and air mixingzone 138 and the fifth annular wall 142 defines the radially innerextremity of the third fuel and air mixing zone 138. A third annular airintake 144 is defined between the upstream ends of the fourth and fifthannular walls 140 and 142 respectively to supply air into the thirdannular fuel and air mixing zones 138.

A plurality of tertiary fuel injectors 146 are provided for each of thetubular combustion chambers 236.

Each fourth and fifth annular wall 140,142 is arranged coaxially aroundthe second portion 42 of the respective annular wall. At the downstreamend of each third fuel and air mixing zone 138, the fourth and fifthannular walls 140 and 142 are secured to the respective fifthfrustoconical portion 132, and each frustoconical portion 132 isprovided with a plurality of circumferentially spaced apertures 148which are arranged to direct fuel and air into a tertiary combustionzone 150, in the tubular combustion chambers 236, in a downstreamdirection towards the axis of the tubular combustion chambers 236. Theapertures 148 may be circular or slots.

In operation primary air A flows through the first air intake 82 andthrough the first and second radial flow swirlers. The lip 78 directsthe primary air into the first fuel and air mixing zone, or primarycombustion zone, 64. The flows of air from the first and second radialflow swirlers are in opposite directions to improve mixing turbulence.

The pilot injectors 86 only are used at low power settings, that is lessthan about 40% power. They inject the gas or pre-evaporated liquid fuelat a narrow angle only into the primary air which has passed through thesecond radial flow swirlers to create a locally fuel rich mixture on theaxes of the tubular combustion chambers 236. Diffusion causes the fuelto mix with the primary air in the vortex B. Vortex C remains an aironly region. Thus a locally rich mixture is created on the combustionchamber 236 centreline which sustains combustion in the primarycombustion zone 64.

The primary fuel injectors 88 are not used, during low power operationand thus only primary air exits from the downstream end of the passages72 and 80 formed between the respective vanes 70 and 74 of the first andsecond swirler assemblies.

At high power settings, at or greater than about 40% power, the pilotinjectors 86 are not used, and all the fuel supplied into the combustionchamber 236 is supplied from the primary and secondary injectors 88 and104 respectively or from the primary, secondary and tertiary injectors88,104 and 146 respectively.

At high power settings, at or greater than about 40% power, the primaryfuel injectors 88 inject gas, or pre-evaporated liquid fuel, into thepassages 72 and 80 formed between the respective vanes 70 and 74 of thefirst and second swirler assemblies. Simultaneously the secondary fuelinjectors 104 inject gas, or pre-evaporated liquid fuel, into the secondfuel and air mixing zone 94 to mix with the secondary air entering thesecond fuel and air mixing zone 94 through the second annular intake102.

The first and second radial flow swirlers direct the fuel and airmixture towards the centreline of the tubular combustion chambers 236before it is turned so that it flows parallel to the centreline of thecombustion chamber 236. The fuel is entrained into both vortex B andvortex C which have opposite swirl, and the shear layer D improvesmixing turbulence.

Secondary air E flows through the second air intake 102 into thesecondary air and fuel mixing zone 94. The secondary air and fuel ismixed as it flows axially downstream through the second fuel and airmixing zone 94. The resulting fuel and air mixture formed in thesecondary air and fuel mixing zone 94 is injected through the apertures106 into the second portion 42 of the tubular combustion chamber 236where secondary combustion occurs in the second combustion zone 112.

The reduction in diameter at the downstream end of the first portion 38,provided by the frustoconical portion 120 and the throat 122 allows asuitable margin between the flame temperature in the primary combustionzone 64 and the temperature at which the flame is extinguished withvariations in power setting to prevent the flame in the primarycombustion zone 64 being extinguished. This enhances the recirculationof hot combustion products into the primary combustion zone 64 toreignite the fuel and mixture.

The reduction in diameter at the downstream end of the second portion40, provided by the frustoconical portion 134 and the throat 136 allowsa suitable margin between the flame temperature in the secondarycombustion zone 112 and the temperature at which the flame isextinguished with variations in power setting to prevent the flame inthe secondary combustion zone 112 being extinguished. This enhances therecirculation of hot combustion products into the secondary combustionzone 112 to reignite the fuel and air mixture, by producingrecirculation zones J.

If the combustion chambers 236 are operated at low ambient airtemperatures, in the range of -60° C. to -30° C., the primary andsecondary fuel injectors 88 and 104 respectively supply fuel into theprimary and secondary combustion zones 64 and 112 respectively for powersettings between 40% and 100% power. The tertiary fuel injectors 146 donot supply fuel into the tertiary combustion zone 150 at low ambient airtemperatures at any power setting. At low ambient air temperatures theamount of fuel supplied to the primary injectors 88 is increased tomaintain the temperature in the primary combustion zone 64 at 1800° K.This is important to ensure optimum combustion for NOx reduction, and tomaintain a high enough temperature in the secondary combustion zone 112for combustion to continue in the secondary combustion zone 112.

If the combustion chambers 236 are operated at high ambient airtemperatures, in the region of +30° C. and above, the primary andsecondary fuel injectors 88 and 104 respectively supply fuel into theprimary and secondary combustion zones 64 and 112 respectively for lowerpower settings between 40% and a predetermined power setting. At highambient air temperatures and high power settings between thepredetermined power setting and 100% power, the primary, secondary andtertiary fuel injectors 88,104 and 146 respectively supply fuel into theprimary, secondary and tertiary combustion zones 64,112 and 150respectively.

As the ambient air temperature is reduced from the high ambient airtemperature, the minimum power setting at which the primary, secondaryand tertiary fuel injectors 88,104 and 146 respectively supply fuel intothe primary, secondary and tertiary combustion zones 64,112 and 150respectively increases from the predetermined power setting at highambient air temperature operation. At low ambient air temperatures, asmentioned previously, the tertiary fuel injectors 146 are not suppliedwith fuel at any power setting.

At high power and high ambient air temperatures, the temperature in thefirst fuel and air mixing zone 64 is maintained at about 1800° K., andthe temperature in the second combustion zone 112 is maintained at about1740° K. and the temperature in the tertiary combustion zone 150 isvaried between 1550° K. and 1800° K. When the temperature in thetertiary combustion zone 150 falls below 1550° K., the tertiary fuelinjectors 146 do not supply fuel to the tertiary combustion zone 150 andthe amount of fuel supplied by the secondary fuel injectors 104 into thesecondary combustion zone 112 is increased to increase its temperatureto 1850° K. The system then acts as a two staged combustor.

The combination of the secondary fuel and air mixing zone 94 andsecondary combustion zone 112 together with the tertiary fuel and airmixing zone 138 and tertiary combustion zone 150 allows reducedemissions of NOx to be achieved at all power settings between 40% and100% power over a wide range of pressure ratios and velocity profileswithout the need for variable geometry air intakes for the combustionchambers 236.

The industrial gas turbine engine will be provided with a control systemwhich controls the fuel supplied to the pilot, primary and secondaryinjectors in accordance with the power demanded for the combustionchambers shown in FIGS. 2 to 5 and 7,

The industrial gas turbine engine will be provided with a control systemwhich controls the fuel supplied to the pilot, primary, secondary andtertiary injectors in accordance with the power demanded and the ambientair temperature for the combustion chamber shown in FIG. 8.

We claim:
 1. A gas turbine engine combustion chamber having alongitudinal axis and comprising first air intake means, primary fuelinjector means and a first fuel and air mixing zone, said first fuel andair mixing zone being defined by at least one annular wall having anupstream end and an upstream wall connected to said upstream end of saidannular wall, said annular wall having a longitudinal axis extendingcoaxially with said longitudinal axis of said combustion chamber atleast partly along said axes, said upstream wall having at least oneaperture, said first air intake means comprising at least one first flowswirler and at least one second flow swirler for introducing first airinto said first fuel and air mixing zone through said aperture in saidupstream wall, said first flow swirler and said second flow swirlerbeing disposed at least partly radially with respect to saidlongitudinal axis and upstream of said annular wall along said axis withsaid first flow swirler being located closer to said end wall than saidsecond flow swirler, said first flow swirler having vanes to swirl airin one direction, said second flow swirler having vanes to swirl air ina direction generally opposite to said one direction, said vanes of eachsaid flow swirler defining passages therebetween, said primary fuelinjector means being located to supply fuel into at least one of saidpassages between said vanes of said first flow swirler and said vanes ofsaid second flow swirler.
 2. The invention as claimed in claim 1,wherein said combustion chamber includes at least one pilot fuelinjector aligned with said aperture in said end wall to supply fuelthrough said aperture into said first fuel and air mixing zone.
 3. Theinvention as claimed in claim 1, wherein said primary fuel injectormeans is located to supply fuel into each of said passages between saidvanes of said first flow swirler.
 4. The invention as claimed in claim1, in which said primary fuel injector means is located to supply fuelinto all the passages defined between said vanes of said second flowswirler.
 5. The invention as claimed in claim 1, in which said passageshave radially outer regions and said primary fuel injector means islocated to supply fuel to said radially outer regions.
 6. The inventionas claimed in claim 1, wherein said primary fuel injector meanscomprises a hollow cylindrical member located to extend axially withrespect to said combustion chamber, said cylindrical member having aplurality of apertures spaced apart along said cylindrical member toinject fuel into said passage.
 7. The invention as claimed in claim 6,wherein said apertures are positioned to direct the fuel radiallyinwardly relative to said axis of said combustion chamber.
 8. Theinvention as claimed in claim 1, in which said combustion chamber istubular and has a single aperture in said upstream wall.
 9. Theinvention as claimed in claim 2, further comprising secondary air intakemeans, secondary fuel injector means and a secondary fuel and air mixingzone, said secondary fuel and air mixing zone being annular andsurrounding said first fuel and air mixing zone, said secondary fuel andair mixing zone having a radially outer extremity defined by a secondannular wall, said secondary fuel injector means being located to supplyfuel into said upstream end of said secondary fuel and air mixing zone,said secondary fuel and air mixing zone having a downstream end in fluidflow communication with a secondary combustion zone provided in saidcombustion chamber downstream of said first fuel and air mixing zone.10. The invention as claimed in claim 9, wherein said annular wall has afirst portion defining said first fuel and air mixing zone, a secondportion of increased diameter downstream of said first portion anddefining said secondary combustion zone, and a third frusto-conicalportion interconnecting the first and second portions.
 11. The inventionas claimed in claim 10, wherein said downstream end of said firstportion of said second annular wall reduces in diameter to a throat. 12.The invention as claimed in claim 9, in which said secondary air intakemeans is downstream of said first air intake means.
 13. The invention asclaimed in claim 9, wherein said secondary fuel and air mixing zone isdefined at its radially inner extremity by a third annular wall.
 14. Theinvention as claimed in claim 10, wherein said third frusto-conicalportion has a plurality of equally circumferentially spaced aperturesfor directing a secondary fuel and air mixture from said secondary fueland air mixing zone as a plurality of jets in a downstream directiontowards said axis of said combustion chamber.
 15. The invention asclaimed in claim 14, in which said apertures are slots.
 16. Theinvention as claimed in claim 10, in which said second annular wall hasa downstream end which is secured to said third frusto-conical portionof said second annular wall.
 17. The invention as claimed in claim 13,wherein said combustion chamber has means for supplying cooling air toan annular chamber defined between said annular wall and said thirdannular wall.
 18. The invention as claimed in claim 9, wherein saidsecondary fuel injector means comprises a plurality ofequi-circumferentially spaced injectors.
 19. The invention as claimed inclaim 9, further comprising tertiary air intake means, tertiary fuelinjector means and a tertiary fuel and air mixing zone, said tertiaryfuel and air mixing zone being annular in shape and surrounding saidsecondary combustion zone, said tertiary fuel and air mixing zone beingdefined at its radially outer extremity by a fourth annular wall, saidtertiary fuel injector means being located to supply fuel into theupstream end of said tertiary fuel and air mixing zone, said tertiaryfuel and air mixing zone being in fluid flow communication at itsdownstream end with a tertiary combustion zone provided in saidcombustion chamber downstream of said secondary combustion zone.
 20. Theinvention as claimed in claim 19 wherein said annular wall has a fourthportion of larger diameter than said second portion downstream of saidsecond portion and defining the tertiary combustion zone, a fifthfrusto-conical portion interconnecting said second and fourth portions.21. The invention as claimed in claim 19, wherein said downstream end ofsaid first portion of said second annular wall reduces in diameter to athroat.
 22. The invention as claimed in claim 19, in which said tertiaryair intake means is downstream of said second air intake means.
 23. Theinvention as claimed in claim 19, wherein said tertiary fuel and airmixing zone is defined at its radially inner extremity by a fifthannular wall.
 24. The invention as claimed in claim 20, wherein saidfifth frusto-conical portion has a plurality of equi-circumferentiallyspaced apertures for directing a tertiary fuel and air mixture from saidtertiary fuel and air mixing zone as a plurality of jets in a downstreamdirection towards said axis of said combustion chamber.
 25. Theinvention as claimed in claim 24, in which said apertures are slots. 26.The invention as claimed in claim 20, in which said fourth annular wallhas a downstream end which is secured to said fifth frusto-conicalportion of said annular wall.
 27. The invention as claimed in claim 19,wherein said secondary fuel injector means comprises a plurality ofequi-circumferentially spaced injectors.
 28. A method as claimed inclaim 27 in which the predetermined output level is 35-40% power.
 29. Amethod as claimed in claim 28 in which the proportion of fuel suppliedfrom the primary fuel injector means varies from 75% to 50% of the totalfuel supplied into the combustion chamber from 40% to 100% output powerlevel.
 30. A method of operating a gas turbine engine combustion chamberof the type having a first fuel intake means, primary fuel injectormeans and a first fuel and air mixing zone, the said zone being definedby at least one annular wall and an upstream wall with said upstreamwall being connected to said upstream end of said annular wall, saidannular wall having a longitudinal axis extending coaxially with saidlongitudinal axis of said combustion chamber, said upstream wall havingat least one aperture, said first air intake means comprising at leastone first flow swirler and at least one second flow swirler forintroducing first air into said first fuel and air mixing zone throughsaid aperture in said upstream wall, said first flow swirler and saidsecond flow swirler being disposed at least partly radially with respectto said longitudinal axis and upstream of said annular wall along saidaxis with said first flow swirler being located closer to said end wallthan said second flow swirler, said first flow swirler having vanes toswirl air in one direction, said second flow swirler having vanes toswirl air in a direction generally opposite to said one direction, saidvanes of each said flow swirler defining passages therebetween, saidprimary fuel injector means being located to supply fuel into at leastone of said passages between said vanes of said first flow swirler andsaid vanes of said second flow swirler, said combustion chamber furtherincluding a pilot fuel injector, the method comprising:supplying fuelfrom the pilot fuel injector only into said first fuel and air mixingzone from the start of operation of the gas turbine engine until apredetermined output power level is obtained, supplying fuel from saidprimary fuel injector means into at least one of the passages definedbetween said vanes of said first flow swirler and into at least one ofsaid passages defined between said vanes of said second flow swirler toflow into said first fuel and air mixing zone for output power levelgreater than the predetermined level, and simultaneously supplying fuelfrom said secondary fuel injector means into said secondary fuel and airmixing zone to flow into said secondary combustion zone provided in theinterior of said combustion chamber downstream of said first fuel andair mixing zone.
 31. A method as claimed in claim 30 in which thepredetermined output power level is 35 to 40% power.
 32. A method ofoperating a gas turbine engine combustion chamber of the type having afirst fuel intake means, primary fuel injector means and a first fueland air mixing zone, the said zone being defined by at least one annularwall and an upstream wall with said upstream wall being connected tosaid upstream end of said annular wall, said annular wall having alongitudinal axis extending coaxially with said longitudinal axis ofsaid combustion chamber, said upstream wall having at least oneaperture, said first air intake means comprising at least one first flowswirler and at least one second flow swirler for introducing first airinto said first fuel and air mixing zone through said aperture in saidupstream wall, said first flow swirler and said second flow swirlerbeing disposed at least partly radially with respect to saidlongitudinal axis and upstream of said annular wall along said axis withsaid first flow swirler being located closer to said end wall than saidsecond flow swirler, said first flow swirler having vanes to swirl airin one direction, said second flow swirler having vanes to swirl air ina direction generally opposite to said one direction, said vanes of eachsaid flow swirler defining passages therebetween, said primary fuelinjector means being located to supply fuel into at least one of saidpassages between said vanes of said first flow swirler and said vanes ofsaid second flow swirler, said combustion chamber further including apilot fuel injector, the method comprising:supplying fuel from saidpilot fuel injector only into said first fuel and air mixing zone fromthe start of operation of the gas turbine engine until a predeterminedoutput power level is obtained, supplying fuel from said primary fuelinjector means into at least one of the passages defined between thevanes of said first flow swirler and into at least one of the passagesdefined between said vanes of said second flow swirler to flow into thefirst fuel and air mixing zone for output power levels greater than thepredetermined level, and simultaneously supplying fuel into thesecondary fuel and air mixing zone to flow into the secondary combustionzone provided in the interior of a combustion chamber downstream of thefirst fuel and air mixing zone, supplying fuel into the tertiary fueland air mixing zone to flow into the said tertiary combustion zoneprovided in the interior of said combustion chamber downstream of saidsecondary combustion zone for output power level greater than a secondpredetermined level and for ambient air temperature greater than apredetermined temperature.
 33. The method as claimed in claim 32introducing the step of injecting gas fuel with said primary fuelinjector means.
 34. The method as claimed in claim 32 including the stepof injecting evaporated liquid fuel with said primary fuel injectormeans.
 35. A gas turbine engine combustion chamber having a longitudinalaxis and comprising first air intake means, primary fuel injector meansand a first fuel and air mixing zone, said first fuel and air mixingzone being defined by at least one annular wall having an upstream endand an upstream wall connected to said upstream end of said annularwall, said annular wall having a longitudinal axis extending coaxiallywith said longitudinal axis of said combustion chamber at least partlyalong said axes, said upstream wall having at least one aperture, saidfirst air intake means comprising at least one first flow swirler and atleast one second flow swirler for introducing first air into said firstfuel and air mixing zone through said aperture in said upstream wall,said first flow swirler and said second flow swirler being disposed atleast partly radially with respect to said longitudinal axis andupstream of said annular wall along said axis with said first flowswirler being located closer to said end wall than said second flowswirler, said first flow swirler having vanes to swirl air in onedirection, said second flow swirler having vanes to swirl air in adirection generally opposite to said one direction, said vanes of eachsaid flow swirler defining passages therebetween, said primary fuelinjector means being located to supply fuel into at least one of saidpassages between said vanes of said first flow swirler and said vanes ofsaid second flow swirler,said combustion chamber including at least onepilot fuel injector aligned with said aperture in said end wall tosupply fuel through said aperture into said first fuel and air mixingzone, said combustion chamber further comprising secondary air intakemeans, secondary fuel injector means and a secondary fuel and air mixingzone, said secondary fuel and air mixing zone being annular andsurrounding said first fuel and air mixing zone, said secondary fuel andair mixing zone having a radially outer extremity defined by a secondannular wall, said secondary fuel injector means being located to supplyfuel into said upstream end of said secondary fuel and air mixing zone,said secondary fuel and air mixing zone having a downstream end in fluidflow communication with a secondary combustion zone provided in saidcombustion chamber downstream of said first fuel and air mixing zone,said annular wall having a first portion defining said first fuel andair mixing zone, a second portion of increased diameter downstream ofsaid first portion and defining said secondary combustion zone, and athird frusto-conical portion interconnecting the first and secondportions, said downstream end of said first portion of said secondannular wall reducing in diameter to a throat.
 36. A gas turbine enginecombustion chamber having a longitudinal axis and comprising first airintake means, primary fuel injector means and a first fuel and airmixing zone, said first fuel and air mixing zone being defined by atleast one annular wall having an upstream end and an upstream wallconnected to said upstream end of said annular wall, said annular wallhaving a longitudinal axis extending coaxially with said longitudinalaxis of said combustion chamber at least partly along said axes, saidupstream wall having at least one aperture, said first air intake meanscomprising at least one first flow swirler and at least one second flowswirler for introducing first air into said first fuel and air mixingzone through said aperture in said upstream wall, said first flowswirler and said second flow swirler being disposed at least partlyradially with respect to said longitudinal axis and upstream of saidannular wall along said axis with said first flow swirler being locatedcloser to said end wall than said second flow swirler, said first flowswirler having vanes to swirl air in one direction, said second flowswirler having vanes to swirl air in a direction generally opposite tosaid one direction, said vanes of each said flow swirler definingpassages therebetween, said primary fuel injector means being located tosupply fuel into at least one of said passages between said vanes ofsaid first flow swirler and said vanes of said second flow swirler,saidcombustion chamber including at least one pilot fuel injector alignedwith said aperture in said end wall to supply fuel through said apertureinto said first fuel and air mixing zone, said combustion chamberfurther comprising secondary air intake means, secondary fuel injectormeans and a secondary fuel and air mixing zone, said secondary fuel andair mixing zone being annular and surrounding said first fuel and airmixing zone, said secondary fuel and air mixing zone having a radiallyouter extremity defined by a second annular wall, said secondary fuelinjector means being located to supply fuel into said upstream end ofsaid secondary fuel and air mixing zone, said secondary fuel and airmixing zone having a downstream end in fluid flow communication with asecondary combustion zone provided in said combustion chamber downstreamof said first fuel and air mixing zone, said combustion chamber furthercomprising tertiary air intake means, tertiary fuel injector means and atertiary fuel and air mixing zone, said tertiary fuel and air mixingzone being annular in shape and surrounding said secondary combustionzone, said tertiary fuel and air mixing zone being defined at itsradially outer extremity by a fourth annular wall, said tertiary fuelinjector means being located to supply fuel into the upstream end ofsaid tertiary fuel and air mixing zone, said tertiary fuel and airmixing zone being in fluid flow communication at its downstream end witha tertiary combustion zone provided in said combustion chamberdownstream of said secondary combustion zone, said annular wall having afourth portion of larger diameter than said second portion downstream ofsaid second portion and defining the tertiary combustion zone, a fifthfrusto-conical portion interconnecting said second and fourth portions.37. The invention as claimed in claim 36 wherein the downstream end ofsaid second portion of said annular wall reduces in diameter to athroat.